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Full text of "Research abstracts /National Advisory Committee for Aeronautics"

National Advisory Committee for Aeronautics 



N0.91 



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esearc 



h Abstracts 



OCTOBER 21, 1955 



CURRENT NACA REPORTS 



NACA Rept. 1191 

ON THE DEVELOPMENT OF TURBULENT WAKES 
FROM VORTEX STREETS. Anatol Roshko, Cali- 
fornia Institute of Technology. 1954. ii, 25p. 
diagrs., photos., 3 tabs. (NACA Rept. 1191. 
Formerly TN 2913) 

Wake development behind circular cylinders at 
Reynolds numbers from 40 to 10,000 was investigated 
by hot-wire techniques in a low-speed wind tunnel. 
The Reynolds, number range of periodic vortex 
shedding is divided into two distinct subranges. In 
the stable range, R = 40 to 150, regular vortex 
streets are formed and no turbulent motion develops, 
the vortices decaying by viscous diffusion. The 
range.R = 150 to 300 is a transition region to the 
irregular range in which turbulent velocity fluctua- 
tions accompany the periodic formation of vortices. 
The diffusion is turbulent and the wake becomes 
fully turbulent in 40 to 50 diameters. The turbulence 
is initiated by laminar-turbulent transition in the 
free layers which spring from the separation points 
on the cylinder. An annular vortex street was ob- 
served in the wake of a ring. 



NACA RM E55E18 

PERFORMANCE CHARACTERISTICS OF HEMI- 
SPHERICAL TARGET-TYPE THRUST REVERSERS. 
Fred W. Steffen, Jack G. McArdle, and James W. 
Coats. September 1955. 39p. diagrs., photos., 
tab. (NACA RM E55E18) 

Reverse-thrust performance of hemispherical 
target -type thrust reversers was obtained. The 
value of reverse-thrust ratio was found to be pri- 
marily a function of hemisphere diameter. Slight 
improvements in performance were obtained with 
some size reversers by operating with the exhaust 
nozzle opened slightly. When high values of 
reverse-thrust ratio were obtained, the reversed 
flow attached to the boattail and the pressures on the 
boattail fell belovv atmospheric. This pressure re- 
duction amounted to about 20 percent of the resultant 
reverse force. Various boattail shapes had little ef- 
fect on reverse-thrust ratio for most hemisphere 
sizes. 



NACA RM E55G27a 

IDEAL TEMPERATURE RISE DUE TO CONSTANT- 
PRESSURE COMBUSTION OF A JP-4 FUEL. 
S(idney) C. Huntley. September 1955. 53p. 
diagrs., tabs. (NACA RM E55G27a) 

Charts are presented from which ideal temperature 
rise or the ideal quantity of fuel required to obtain a 
specified combustion temperature may be obtained. 
The charts are applicable only to a fuel having a 
hydrogen-carbon mass ratio of 0.168 (CH2) and in- 
clude a range of fuel-air ratios from to 1.2 frac- 
tions of stoichiometric fuel-air ratio >vith dissocia- 
tion taken into account, inlet-air temperatures from 
400° to 1600° R, and combustion pressures from' 
1/16 to 64 atmospheres 



NACA RM L55G21 



NOTE ON HOVERING TURNS WITH TANDEM 
HELICOPTERS. John P. Reeder and Robert J. 
Tapscott. September 1955. 5p. photo. (NACA 
RM L55G21) 

The source of an appreciable pitching-moment dif- 
ference between left and right hovering turns for a 
tandem helicopter is described. The difference in 
pitching moment results from the difference in ro- 
tational speed of the counterrotating rotors with 
respect to the air while the helicopter is turning. 



NACA TM 1377 

THE THEORIES OF TURBULENCE. (Les Theories 
de la Turbulence). L. Agostini and J. Bass. 
October 1955. 163p. diagrs. (NACA TM 1377. 
Trans, from Ministere de l'Air, Publications 
Scientifiques et Techniques 237, 1950) 

The report includes a discussion of the kinematics 
of statistical mediums, particularly those which are 
isotropic. A mathematical study is made of the ap- 
plications of Navier's equations to turbulent motion. 
Physical theories involving similarity are dealt with. 
Review is made of much of the work in turbulence. 
The theoretical discussions are illustrated by some 
correlation and spectrum curves based on measure- 
ments taken in the wind tunnel at the laboratory of 
the mechanics of the atmosphere at Marseille. 



U.S. DEPOSITORY } 



•AVAILABLE ON LOAN ONLY, 

ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 1512 H ST., NW„ WASHINGTON 25, D C ., CITING CODE NUMBER ABOVE EACH TITLE; 

THE REPORT TITLE AND AUTHOR. 



2.% /3CX? 



NACA TM 1389 

OPTIMUM FLIGHT PATHS OF TURBOJET AIR- 
CRAFT. (Traiettorie Ottime Di Volo Degli Aero- 
plani Azionati Da Turboreattori). Angelo Miele. 
September 1955. 47p. diagrs. , tabs. (NACA TM 
1389. Trans, from L'Aerotecnica, v. 32, no. 4, 
1952, p. 206-219) 

The climb of turbojet aircraft is analyzed and dis- 
cussed including the effects of tangential accelera- 
tions. Three flight performances are examined: 
minimum time of climb, climb with minimum fuel 
consumption, and steepest climb. Diagrams for 
quick calculation of the optimum Mach numbers and 
the effect of acceleration on the rate of climb in 
tropospheric and stratospheric flight are given. 



NACA TM 1396 

FROM LINEAR MECHANICS TO NONLINEAR 
MECHANICS. (De la mecanique lineaire a la 
mecanique non lineaire). Julien Loeb. October 
1955. 18p. diagrs.^ ^NACA TM 1396. Trans, 
from Annales des Telecommunications, v. 5, no. 2, 
Feb., 1950, p. 65-71) 

Consideration is first given to the technique used in 
telecommunication where a nonlinear system (the 
modulator) results in a linear transposition of a 
signal. It is then shown that a similar method per- 
mits linearization of electromechanical devices or 
nonlinear mechanical devices. A sweep function 
plays the same role as the carrier wave in radio- 
electricity. The linearizations of certain non- 
linear functionals are presented. 



NACA TN 3463 

INVESTIGATION OF THE VIBRATIONS OF A HOL- 
LOW THIN-WALLED RECTANGULAR BEAM. 
Eldon E. Kordes and Edwin T. Kruszewski. October 
1955. 24p. diagrs., photos., 2 tabs. (NACA 
TN 3463) 

Experimental modes and frequencies of an unstif- 
fened hollow beam of rectangular cross section are 
presented, and comparisons are made between ex- 
perimental and theoretical frequencies. Theories 
based on rigid cross sections were found to be suf- 
ficiently accurate to predict the frequencies of only 
the lovvest three bending modes. For the higher 
bending modes and all the torsional modes, it was 
necessary to include the effects of cross-sectional 
distortions in the calculations. 



NACA TN 3464 

INFLUENCE OF SHEAR DEFORMATION ON THE 
CROSS SECTION ON TORSIONAL FREQUENCIES 
OF BOX BEAMS. Edwin T. Kruszewski and William 
W. Davenport. October 1955. 23p. diagrs. (NACA 
TN 3464) 

An exact analysis has been carried out on the tor- 
sional vibrations of a four-flange box beam with 



NACA 

RESEARCH ABSTRACTS NO. 91 



cross sections which can change shape because the 
stiffness of the bulkheads is finite. The effect of 
shear deformation of the cross section on the tor- 
sional frequencies is illustrated by numerical cal- 
culations. An approximate method for quickly 
estimating the effects of bulkhead shear stiffness on 
the torsional frequencies of box beams has been 
devised. 



NACA TN 3475 

AN ANALYSIS OF ACCELERATION, AIRSPEED, 
AND GUST-VELOCITY DATA FROM ONE TYPE OF 
FOUR-ENGINE TRANSPORT AIRPLANE OPERATED 
OVER TWO DOMESTIC ROUTES. Martin R. Copp 
and Thomas L. Coleman. October 1955. 29p. 

diagrs., tabs. (NACA TN 3475) 

Time-history data obtained by the NACA VGH re- 
corder from one type of four-engine commercial 
transport airplane during operations on two domestic 
routes indicated that the number of gust accelera- 
tions experienced per mile of flight by the two opera- 
tions differed by a factor of roughly 3. The number 
of gusts per mile of flight differed by a factor of 
roughly 4. A general decrease in the frequency of 
occurrence of gust velocities with increasing altitude 
was noted for both operations. For acceleration 
values above 0.8g, maneuver accelerations formed a 
substantial part of the total-flight load histories. A 
comparison of the average overall airspeeds and the 
corresponding average airspeeds in rough air (with 
accelerations equal to or greater than 0.3g) indicated 
very little slowdown by either operation upon en- 
countering turbulence. 



NACA TN 3481 

WIND-TUNNEL INVESTIGATION AT LOW SPEED 
OF EFFECT OF SIZE AND POSITION OF CLOSED 
AIR DUCTS ON STATIC LONGITUDINAL AND 
STATIC LATERAL STABILITY CHARACTERISTICS 
OF UNSWEPT -MID WING MODELS HAVING WINGS 
OF ASPECT RATIO 2, 4, AND 6. Byron M. Jaquet 
and James L. Williams. September 1955. 45p. 
diagrs., photos., tabs. (NACA TN 3481) 

Results are presented of tests made at a Mach num- 
ber of 0.13 in the Langley stability tunnel to deter- 
mine the effects of closed wing-root air ducts on the 
static longitudinal and static lateral stability charac- 
teristics of models having unswept wing and tail sur- 
faces with wings of aspect ratio 2, 4, and 6. In 
addition, for model configurations employing the 
wing of aspect ratio 2 the effects of top and bottom 
fuselage ducts on the static longitudinal and static 
lateral characteristics were determined. The effect 
of the wing-root ducts on the aerodynamic hystere- 
sis in sideslip of the model employing the wing of 
aspect ratio 2 was also determined. 



NACA 

RESEARCH ABSTRACTS NO. 91 



NACA TN 3485 

AN APPROXIMATE SOLUTION FOR AXIALLY 
SYMMETRIC FLOW OVER A CONE WITH AN AT- 
TACHED SHOCK WAVE. Richard A. Hord. 
October 1955. 32p. diagrs. (NACA TN 3485) 

It is shown that the streamlines in an angular 
neighborhood of the surface of an unyawed circular 
cone with an attached shock wave are, to a first ap- 
proximation, portions of hyperbolas. This fact is 
used as a basis for the development of an approxi- 
mate solution in which shock-wave orientation and 
flow field behind the shock wave are given explicitly 
in terms of free-stream Mach number, vertex angle 
of the body cone, and the ratio of specific heats of 
the gas. The approximate solution is compared 
with other approximate solutions for the cone. 



NACA TN 3497 

SUMMARY OF RESULTS OF A WIND-TUNNEL 
INVESTIGATION OF NINE RELATED HORIZONTAL 
TAILS. Jules B. Dods, Jr. and Bruce E. Tinling. 
July 1955. 105p. diagrs., 2 tabs. (NACA TN 3497. 
Formerly RM A51G31a) 

A compilation of data is presented for models of 
nine related horizontal tails. The majority of the 
results were obtained at a Mach number of approxi- 
mately 0.20. Three of the models were tested 
throughout the subsonic Mach number range to a 
maximum of 0.94. The Reynolds number range was 
from 2 to 4 million. The models had aspect ratios 
from 2 to 6, angles of sweepback from 5. 7° to 45°, 
and had 30-percent-chord, sealed, plain flaps. The 
lift coefficient, hinge-moment coefficient, and pres- 
sure coefficients across the elevator nose seal are 
presented. The effects of sweepback, aspect ratio, 
and Mach number on the lift and hinge-moment 
parameters are summarized. Comparisons of the 
experimental results with theoretical calculations 
are presented. 



NACA TN 3519 

VISUALIZATION STUDY OF SECONDARY FLOWS 
IN TURBINE ROTOR TIP REGIONS. Hubert W. 
Allen and Milton G. Kofskey. September 1955. 
33p. diagrs., photos., tab. (NACA TN 3519) 

A low-speed visualization study of turbine rotor tip 
secondary flows was made. Results include qualita- 
tive information on tip-clearance flow, cross- 
passage flow, and scraping flow, and on a range of 
rotor speeds for which a transition condition appear- 
ed with minimum flow disturbance. Rotor speed re- 
quired for this transition flow condition depended 
upon blade configuration, angle of incidence, tip 
clearance, and air-flow rate. Results should aid in 
extending the study to higher airspeeds. 



NACA TN 3532 

LOW-SPEED STATIC LATERAL AND ROLLING 
STABILITY CHARACTERISTICS OF A SERIES OF 
CONFIGURATIONS COMPOSED OF INTERSECTING 
TRIANGULAR PLAN-FORM SURFACES. David F. 
Thomas, Jr. October 1955. 29p. diagrs., photos. 
(NACA TN 3532) 

The static lateral and rolling stability derivatives 
of a series of cruciform, inverted T-, V-, and Y- 
configurations composed of low-aspect-ratio trian- 
gular surfaces have been obtained at low speed in 
the 6-foot-diameter rolling-flow test section of the 
Langley stability tunnel. These derivatives are 
presented as functions of the geometry of the models, 
and for two configurations (a planar wing and an in- 
verted T), as functions of angle of attack. Where 
possible, comparisons have been made to indicate 
the extent of agreement between experiment and 
existing theory. In general, the sideslip deriva- 
tives showed better agreement between theory and 
experiment than the rolling derivatives. 



NACA TN 3533 

THE PROPER COMBINATION OF LIFT LOADING 
FOR LEAST DRAG ON A SUPERSONIC WING. 
Frederick C. Grant. October 1955. 21p. diagrs. 
(NACA TN 3533) 

Lagrange's method of undetermined multipliers is 
applied to the problem of properly combining lift 
loadings for the least drag at a given lift on super- 
sonic wings. The interference drag between the 
optimum loading and any loading at the same lift 
coefficient is found to be constant on a given plan 
form. This is an integral form of a criterion 
established by Robert T. Jones for optimum load- 
ings. The best combination of four loadings on a 
delta wing with subsonic leading edges is calculated 
for several Mach numbers as a numerical example. 



NACA TN 3535 

FLIGHT INVESTIGATION OF THE SURFACE- 
PRESSURE DISTRIBUTION AND THE FLOW FIELD 
AROUND A CONICAL AND TWO SPHERICAL NON- 
ROTATING FULL-SCALE PROPELLER SPINNERS. 
Jerome B. Hammack, Milton L. Windier, and 
Elwood F. Scheithauer. September 1955. 36p. 
diagrs. , photos. (NACA TN 3535) 

The surface-pressure distribution and the flow field 
around a conical and two spherical nonrotating full- 
scale propeller spinners were determined in flight 
at Mach numbers of 0. 70 to 0.96. The local-surface 
Mach numbers between the cone-sphere tangency and 
the maximum thickness stations were approximately 
0.4 higher than those at a free-stream Mach number 
of 0.95. The departure from free-stream conditions 
of the larger spherical spinner extends beyond the 
1. 3-spinner-radius station; whereas, with the 
smaller spherical spinner, free-stream conditions 
were reached at the 1. 3-spinner-radius station. 



NACA TN 3536 

A LIMITED FLIGHT INVESTIGATION OF THE EF- 
FECT OF THREE VORTEX-GENERATOR CONFIGU- 
RATIONS ON THE EFFECTIVENESS OF A PLAIN 
FLAP ON AN UNSWEPT WING. Garland J. Morris 
and Lindsay J(ohn) Lina. September 1955. 20p. 
diagrs. , photos. , tabs. (NACA TN 3536) 

An exploratory flight investigation was made to deter- 
mine the effect of three vortex-generator configura- 
tions on the effectiveness of the plain flap of a fighter 
airplane. Tests were made with flaps deflected 19° 
and 45° at several indicated airspeeds in the range 
from stall to 140 miles per hour. No improvement 
in flap effectiveness was obtained with the flaps de- 
flected full down. With flaps deflected 19°, the 
vortex-generator configuration which produced the 
largest increase in lift coefficient had about the same 
effect as increasing the flap deflection to about 27°. 
Another configuration produced some increase in 
lift coefficient with no apparent increase in drag. 



NACA TN 3538 

SUMMARY OF DERIVED GUST VELOCITIES OB- 
TAINED FROM MEASUREMENTS WITHIN THUN- 
DERSTORMS. H(arold) B. Tolefson. October 
1955. 19p. diagrs., tabs. (NACA TN 3538) 

Available measurements of the derived gust veloci- 
ties within thunderstorms are summarized for alti- 
tudes from 5,000 to 34,000 feet. The results indi- 
cate that the intensity of the derived gust velocity is 
essentially constant up to altitudes of 20,000 feet and 
that an approximate 10-percent reduction in the gust 
intensity occurs for altitudes from 20,000 to 30,000 
feet. 



NACA TN 3561 

INTENSITY, SCALE, AND SPECTRA OF TURBU- 
LENCE IN MIXING REGION OF FREE SUBSONIC 
JET. James C. Laurence. September 1955. 58p. 
diagrs., photo., tab. (NACA TN 3561) 

Hot-wire anemometer measurements of the turbu- 
lence parameters were made in a 3.5-inch-diameter 
free jet at exit Mach numbers between 0.2 and 0.7 
and Reynolds numbers (based on jet rad. ) between 
37,500 and 350,000. The results of these measure- 
ments show that (1) the intensity of turbulence is a 
max. at a distance of approximately 1 jet rad. from 
the jet center line and decreases with increasing 
Mach and/or Reynolds number, and (2) the lateral 
and longitudinal scales of turbulence are nearly in- 
dependent of Mach and/or Reynolds number and vary 
proportionally with distance from the jet nozzle. 
The lateral scale is much smaller than the longi- 
tudinal and does not vary with distance from the 
center line, while the longitudinal scale is a max. 
at a distance from the center line of about 0.7 to 0.8 
of the jet rad. 



NACA 

RESEARCH ABSTRACTS NO. 91 



NACA TN 3568 

AVERAGING OF PERIODIC PRESSURE PULSA- 
TIONS BY A TOTAL-PRESSURE PROBE. R. C. 
Johnson. October 1955. 30p. diagrs., photo., 
tabs. (NACA TN 3568) 

Information is presented on the average pressure 
indicated by a total -pressure probe subjected to a 
stagnation pressure that alternates periodically be- 
tween two constant values. Calculated and experi- 
mental data are in good agreement, and errors are 
reduced when the probe design is such as to ensure 
laminar-flow pulsations in the probe at all times. 
The averaging error is minimized when the inside 
diameter of the probe entrance tube is made as 
small as possible, and its length as great as possi- 
ble, consistent with an acceptable time lag. 



NACA TN 3570 

AN EXPERIMENTAL COMPARISON OF THE 
LAGRANGIAN AND EULERIAN CORRELATION 
COEFFICIENTS IN HOMOGENEOUS ISOTROPIC 
TURBULENCE. William R. Mickelsen. October 
1955. 42p. diagrs. (NACA TN 3570) 

The Lagrangian and Eulerian correlation coeffi- 
cients were compared in a field of homogeneous, 
isotropic turbulence. The Lagrangian correlation 
coefficient was characterized by diffusion measure- 
ments, and the Eulerian coefficient was measured 
by hot-wire anemometry. The Lagrangian and 
Eulerian correlation coefficients had similar shapes 
connected by a linear relation between their coordi- 
nates. The proportionality factor in the linear rela- 
tion was roughly constant over a range of turbulence 
intensities from 1.8 to 14 feet per second. The lin- 
ear relation permits solution of mixing problems 
from the Eulerian turbulence parameters. 



BRITISH REPORTS 



N-38869* 

Royal Aircraft Establishment (Gt. Brit.) 
A SIMPLIFIED MODEL OF THE INCOMPRESSIBLE 
FLOW PAST TWO-DIMENSIONAL AEROFOILS 
WITH A LONG BUBBLE TYPE OF FLOW SEPARA- 
TION. J. F. Norbury and L. F. Crabtree. 
June 19,55. 17p. diagrs., tabs. (RAE Tech. Note 
Aero 2352) 

This problem may be divided into two parts. The 
first part concerns the external inviscid flow, and 
the other is related to the details of the flow inside 
the viscous region formed by the bubble, the bound- 
ary layer, and the wake. Although much informa- 
tion can be obtained by considering only the external 
stream (as in the hodograph method developed by 
Maskell), a complete and unique solution for the 
pressure distribution can only be found by attacking 
the second part of the problem as well. This has 
been attempted here and it is shown how the pres- 
sure recovery ratio at the end of a long bubble may 
be found by the analysis of a simplified model of the 
flow; the complete pressure distribution may then 
be uniquely determined. 



NACA 

RESEARCH ABSTRACTS NO. 91 



MISCELLANEOUS 



N-39013* 

AIRCRAFT STRUCTURES RESEARCH AT ELEVAT- 
ED TEMPERATURES. John E. Duberg. (Present- 
ed to Structures and Materials Panel of Advisory 
Group for Aeronautical Research and Development 
(NATO) London, England, September 5-9, 1955). 
36p. diagrs. , photos. 

A review is made of the test techniques that have 
been developed and used by the NACA for experimen- 
tal research in aircraft structures at elevated temp- 
eratures. Some experimental results are presented. 
Remarks are included on the problem of model scal- 
ing for testing of structures at high temperatures. 



UNPUBLISHED PAPERS 



N-37996* 

THE BUCKLING OF PLATES AND BARS IN THE 
PLASTIC RANGE. I. THEORY. (Over het 
knitvraagstuk in het plastische gebied bij staven en 
platen. I. Theorie). J. F. Besseling. July 1955. 
97p. diagrs., tabs. (Trans, from Nationaal 
Luchtvaartlaboratorium, Amsterdam, S.407, Oct. 14, 
1952) 

The topics treated include the plastic buckling of 
bars and plates; application of plasticity theories to 
'the determination of stiffness quantities needed for 
computing the buckling; and the effect of plasticity 
on the buckling load for a number of fundamental 
buckling cases when the buckling stress condition is 
reached through a uniform increase of loading. This 
report is limited to the theoretical side of the buck- 
ling problem and gives the basis for comparing 
theoretical and experimental results. 



DECLASSIFIED NACA REPORTS 



NACA RM A8I17 

INVESTIGATION OF A THIN WING OF ASPECT 
RATIO 4 IN THE AMES 12-FOOT PRESSURE WIND 
TUNNEL. Ill - THE EFFECTIVENESS OF A 
CONSTANT-CHORD AILERON. Ben H. Johnson, Jr. 
and Fred A. Demele. November 19, 1948. 26p. 
diagrs., photo. (NACA RM A8I17) 
(Declassified from Confidential, 9/15/55) 

Presented in the report are results of tests at Mach 
numbers from 0.27 to 0.94 of a thin, unswept wing 



having a modified diamond airfoil section of thickness 
ratio 0.042 and equipped with a constant-chord ailer- 
on. The tests were conducted at a constant Reynolds 
number of 2, 730, 000. The effects of compressibility 
on the aileron effectiveness were negligible at Mach 
numbers up to 0.85, but at higher Mach numbers er- 
ratic effects of compressibility were evident, espe- 
cially at lift coefficients greater than 0.5. 



NACA RM A9I01 

INVESTIGATION OF A THIN WING OF ASPECT 
RATIO 4 IN THE AMES 12-FOOT PRESSURE WIND 
TUNNEL. V - STATIC LONGITUDINAL STABILITY 
AND CONTROL THROUGHOUT THE SUBSONIC 
SPEED RANGE OF A SEMISPAN MODEL OF A 
SUPERSONIC AIRPLANE. Ben H. Johnson, Jr. . 
and Francis W. Rollins. December 8, 1949. 130p. 
diagrs., photos. (NACA RM A9I01) 
(Declassified from Confidential, 9/15/55) 

Presented in this report are results of tests at Mach 
numbers from 0.20 to 0.94 of a model of a hypo- 
thetical supersonic airplane equipped with a thin, 
sharp-edged wing and tail without sweep. The static 
longitudinal-stability characteristics of the model 
have been measured for two different vertical loca- 
tions of the horizontal tail. The longitudinal control 
afforded by an all-movable stabilizer and by an ele- 
vator has been investigated. The downwash at the 
tail has been computed and the dynamic-pressure 
ratio at the tail has been evaluated from pressure 
measurements in the wake of the wing. 



NACA RM A51I07 

THE STATIC LONGITUDINAL CHARACTERISTICS 
AT MACH NUMBERS UP TO 0.95 OF A 
TRIANGULAR-WING CANARD MODEL HAVING A 
TRIANGULAR CONTROL. Jack D. Stephenson and 
Ralph Selan. December 1951. 72p. diagrs., photo. 
(NACA RM A51I07) (Declassified from 
Confidential, 9/15/55) 

Presents and analyzes results of tests to assess the 
longitudinal characteristics of a canard-type model 
having a triangular wing of aspect ratio 2 and NACA 
0008-63 sections. The horizontal canard surface 
had a plan form identical to the wing and an NACA 
0005-63 section. The Mach number range of the 
investigation varied from 0.25 to 0.95 at Reynolds 
numbers of 8 million and 3 million, respectively. 
The model was tested with the horizontal control 
surface at various fixed angles of incidence and with 
the surface unrestrained so that it could pivot about 
an axis at 30 percent of its mean aerodynamic chord. 



NACA RM A52F18 

THE LONGITUDINAL CHARACTERISTICS AT MACH 
NUMBERS UP TO 0.92 OF A CAMBERED AND 
TWISTED WING HAVING 40° OF SWEEPBACK AND 
AN ASPECT RATIO OF 10. George G. Edwards, 
Bruce E. Tinling, and Arthur C. Ackerman. 
September 1952. 71p. diagrs., photos., tab. 
(NACA RM A52F18) (Declassified from 
Confidential, 9/15/55) 

A sweptback wing, in combination with a fuselage, of 
a type considered suitable for long-range, high- 
speed airplanes, has been tested in the Ames 12-foot 
pressure wind tunnel. The wing had 40° of sweep- 
back, an aspect ratio of 10, a taper ratio of 0.4, and 
5° of washout at the tip. The sections normal to the 
sweep reference line had NACA four-digit profiles, 
design lift coefficients of 0.40, and varied in thick- 
ness ratio from 14 percent at the root to 11 percent 
at the tip. The lift, drag, and pitching moment of a 
semispan model were measured at Reynolds num- 
bers from 2,000,000 to 8,000.000 at low Mach num- 
bers, and at Mach numbers from 0.25 to 0.92 at a 
Reynolds number of 2,000,000. The changes in 
boundary-layer flow on the upper surface were 
studied with tufts. 



NACA RM A52I19 

THE LONGITUDINAL CHARACTERISTICS AT MACH 
NUMBERS UP TO 0.9 OF A WING-FUSELAGE-TAIL 
COMBINATION HAVING A WING WITH 40° OF 
SWEEPBACK AND AN ASPECT RATIO OF 10. 
Bruce E. Tinling. December 1952. 41p. diagrs., 
photo., tab. (NACA RM A52I19) (Declassified from 
Confidential, 9/15/55) 

Wind-tunnel tests were made of a wing-fuselage hori- 
zontal tail combination suitable for long-range high- 
speed airplanes. The cambered and twisted wing 
had an aspect ratio of 10, taper ratio of 0.4, and 40° 
sweepback. The all-movable horizontal tail had an 
aspect ratio of 4.5, taper ratio of 0.4, and 40° 
sweepback. Included are data on longitudinal sta- 
bility and control, wing fence and tail effect, and 
isolated horizontal tail data. Data »vere obtained at 
a Reynolds number of 8,000,000 at low Mach num- 
bers, and at a Reynolds number of 2,000,000 at Mach 
numbers up to 0.9. 



NACA RM A52K20 

PRESSURE DISTRIBUTION AT MACH NUMBERS UP 
TO 0.90 ON A CAMBERED AND TWISTED WING 
HAVING 40° OF SWEEPBACK AND AN ASPECT 
RATIO OF 10, INCLUDING THE EFFECTS OF 
FENCES. Frederick W. Boltz and Harry H. 
Shibata. March 1953. 133p. diagrs., photos., tabs. 
(NACA RM A52K20) (Declassified from 
Confidential, 9/15/55) 

Pressure-distribution measurements have been made 
on a semispan model of a cambered and twisted wing, 



NACA 
RESEARCH 



ABSTRACTS NO. 9) 



alone and in combination with a fuselage. The wing 
had 40° of sweepback, an aspect ratio of 10, a taper 
ratio of 0.4, and 5° of washout at the tip. The wing 
thickness distribution in sections normal to the ref- 
erence sweep line was the NACA 4-digit series with 
the maximum thickness varying from 14-percent 
chord at the root to 11-percent chord at the tip. The 
chordwise distributions of pressure coefficient at 
nine semispan stations are presented for Mach num- 
bers of 0.165 and 0.25 at a Reynolds number of 
8,000,000, and for Mach numbers from 0.25 to 0.90 
at a Reynolds number of 2,000,000. Tabulated 
pressure data are presented for the wing without 
fences and with four fences. 



NACA RM A53D06 

THE EFFECTS OF NACELLES AND OF EXTENDED 
SPLIT FLAPS ON THE LONGITUDINAL CHARAC- 
TERISTICS OF A WING-FUSELAGE-TAIL COMBI- 
NATION HAVING A WING WITH 40° OF SWEEP- 
BACK AND AN ASPECT RATIO OF 10. Bruce E. 
Tinling and Armando E. Lopez. June 1953. 47p. 
diagrs., tab. (NACA RM A53D06) (Declassified 
from Confidential, 9/15/55) 

Wind-tunnel tests were made to evaluate the effects 
of nacelles and of extended split flaps on the longitu- 
dinal characteristics of a wing-fuselage-tail combi- 
nation suitable for long-range high-speed airplanes. 
The cambered and twisted wing had an aspect ratio 
of 10, a taper ratio of 0.4, and 40° of sweepback. 
The nacelles were at 25 and 50 percent of the semi- 
span. Data were obtained to evaluate the effects of 
nacelles at Mach numbers up to 0.90 at a Reynolds 
number of 2,000,000. The effects of the flaps were 
evaluated from data obtained at a Reynolds number 
of 4,000,000 and a Mach number of 0.082. A 
limited number of data were also obtained to evalu- 
ate the effects of increasing the tail height. 



NACA RM A53I23 

DOWNWASH BEHIND A TRIANGULAR WING OF 
ASPECT RATIO 3 - TRANSONIC BUMP METHOD. 
John A. Axelson. December 1953. 37p. diagrs., 
photo., tab. (NACA RM A53I23) (Declassified from 
Confidential, 9/15/55) 

Downwash measured by means of an all-movable 
horizontal tail in several different locations is pre- 
sented for a triangular wing having an aspect ratio of 
3 and the NACA 63A006 section. The lift, drag, and 
pitching-moment characteristics of the wing and the 
downwash are presented for angles of attack up to 
28° over a Mach number range from 0.6 to 1.1, 
corresponding to a Reynolds number range from 1.8 
million to 2.4 million. The effects of tail location 
on the tail contribution to the static longitudinal sta- 
bility of the wing-tail combination are discussed. 



NACA 

RESEARCH ABSTRACTS NO. 91 



NACA RM A53I28 

THE RESULTS OF WIND-TUNNEL TESTS AT LOW 
SPEEDS OF A FOUR-ENGINE PROPELLER-DRIVEN 
AIRPLANE CONFIGURATION HAVING A WING WITH 
40° OF SWEEPBACK AND AN ASPECT RATIO OF 
10. George G. Edwards, Donald A. Buell, and 
Jerald K. Dickson. December 1953. 121p. diagrs., 
photo., tabs. (NACA RM A53I28) (Declassified 
from Confidential, 9/15/55) 

The effects of operating propellers on the low-speed 
longitudinal characteristics of a four-engine tractor 
airplane configuration having a sweptback wing have 
been investigated in the Ames 12-foot pressure wind 
tunnel at thrust coefficients up to 0.9 per propeller 
and at Reynolds numbers from 4,000,000 to 
8,000,000. Variations in the model included different 
heights and incidences of the horizontal tail as well 
as tail removed, two arrangements of extended split 
flaps, several propeller-blade angles, and inde- 
pendent as well as simultaneous operation of the in- 
board and outboard propellers. Coefficients of lift, 
longitudinal force, pitching moment, propeller 
thrust, and propeller power are presented in tabular 
form for various values of advance ratio at constant 
angles of attack. Selected portions of the data are 
presented in plotted form for various constant thrust 
coefficients. 



NACA RM A53I29 

THE TRANSONIC CHARACTERISTICS OF 36 SYM- 
METRICAL WINGS OF VARYING TAPER, ASPECT 
RATIO, AND THICKNESS AS DETERMINED BY THE 
TRANSONIC -BUMP TECHNIQUE. Warren H. 
Nelson, Edwin C. Allen, and Walter J. Krumm. 
December 1953. 131p. diagrs., photo. (NACA 
RM A53I29) (Declassified from Confidential, 
9/15/55) 

An investigation was made in the Ames 16-foot high- 
speed wind tunnel, utilizing the transonic-bump tech- 
nique to determine the effects of plan-form taper on 
a series of wings having aspect ratios of 4, 3, and 2, 
and NACA 63A00X sections with thickness-to-chord 
ratios of 8, 6, 4, and 2 percent. The Mach number 
range was 0.6 to 1.1 with a corresponding Reynolds 
number range of about 1.4 million to 2.0 million. 
The results indicate that increasing the taper ratio 
caused only small increases in lift-curve slope 
except for the wings of highest aspect ratio and 
thinnest sections. Increasing taper ratio generally 
increased the overall center-of-pressure travel in 
going from subsonic to supersonic speeds. 



NACA RM E8I01 

EFFECT OF THREE FLAME -HOLDER CONFIGU- 
RATIONS ON SUBSONIC FLIGHT PERFORMANCE 
OF RECTANGULAR RAM JET OVER RANGE OF 
ALTITUDES. Dugald O. Black and Wesley E. 
Messing. November 24, 1948. 28p. diagrs., photo., 
tab. (NACA RM E8I01) 
(Declassified from Confidential, 9/15/55) 

A flight investigation is reported that was conducted 
on a rectangular ram jet incorporating various 



flame-holder configurations over a range of fuel -air 
ratios from 0.017 to 0.120, combustion-chamber- 
inlet velocities from 50 to 125 feet per second, and 
pressure altitudes from 1500 to 28,000 feet. Highest 
combustion efficiencies, which varied from a maxi- 
mum of 82 percent at 1500 feet to 39 percent at 
26,000 feet, were obtained at all altitudes with the 
three-V gutter-type flame holder. However, at any 
given altitude and fuel-air ratio slightly higher net- 
thrust coefficients occurred with the two-V flame 
holder as a result of its lower value of pressure 
drop. 



NACA RM E8I28 

EFFECT OF VARIATION IN FUEL PRESSURE ON 
COMBUSTION PERFORMANCE OF RECTANGULAR 
RAM JET. Wesley E. Messing and Dugald O. Black. 
November 24, 1948. 26p. diagrs., photo., tab. 
(NACA RM E8I28) 
(Declassified from Confidential, 9/15/55) 

Reports effect of variation in fuel pressure on start- 
ing characteristics, minimum blow-out limits, com- 
bustion efficiencies, gas total-temperature ratio, and 
net -thrust coefficient of rectangular ram jet operated 
over range of pressure altitudes from 1500 to 26, 300 
feet, indicated airspeeds from 100 to 200 miles per 
hour, and fuel-air ratios from 0.017 to 0.120. In 
general, increasing the degree of fuel atomization 
and distribution by utilization of small orifice fuel 
nozzles that operated at high fuel pressures resulted 
in higher values of combustion efficiency, gas total- 
temperature ratio, and net-thrust coefficient at a 
given fuel -air ratio. 



NACA RM E50D28 

AERODYNAMIC CHARACTERISTICS OF NACA RM- 
10 MISSILE IN 8- BY 6-FOOT SUPERSONIC WIND 
TUNNEL AT MACH NUMBERS FROM 1.49 TO 1.98. 
H - PRESENTATION AND ANALYSIS OF FORCE 
MEASUREMENTS. Fred T. Esenwein, Leonard J. 
Obery, and Carl F. Schueller. July 21, 1950. 34p. 
diagrs., photo. (NACA RM E50D28) 
(Declassified from Confidential, 9/15/55) 

Experimental investigation of aerodynamic forces 
acting on body of revolution (NACA RM-10 missile) 
with and without stabilizing fins was conducted at 
Mach numbers from 1.49 to 1.98 at angles of attack 
from 0° to 9° and at Reynolds number of approxi- 
mately 30, 000, 000. Comparison of experimental 
lift, drag, and pitching-moment coefficients and 
center-of-pressure location for body alone is made 
with linearized potential theory and a semiempirical 
method. Results indicate that aerodynamic charac- 
teristics were predicted more accurately by semi- 
empirical method than by potential theory. Break- 
down of measured drag coefficients into components 
of friction, pressure, and base-pressure drag is 
presented for body alone at zero angle of attack. 



NACA RM E 5011 9 

AERODYNAMIC CHARACTERISTICS OF NACA 
RM-10 MISSILE IN 8- BY 6-FOOT SUPERSONIC 
WIND TUNNEL AT MACH NUMBERS FROM 1.49 TO 
1.98. m - ANALYSIS OF FORCE DISTRIBUTION 
AT ANGLE OF ATTACK (STABILIZING FINS RE- 
MOVED). Roger W. Luidens and Paul C. Simon. 
December 12, 1950. 26p. diagrs. 
(NACA RM E50I19) 
(Declassified from Confidential 9/15/55) 

Analysis of force distribution on slender pointed 
body of revolution at angle of attack was made utiliz- 
ing pressure-distribution data and balance measure- 
ments obtained in NACA Lewis 8- by 6-foot super- 
sonic tunnel. Comparison of experimental station 
normal force with those predicted by linearized po- 
tential theory (on which radius of body assumed to 
approach zero) shows that inaccurate prediction by 
theory of normal force acting on slender body of rev- 
olution with curved profiles at angle-of-attack re- 
sults from inaccurate prediction of potential flow 
pressure distribution due to angle of attack, and 
from neglecting effect of cross-flow separation. In- 
crease in total axial force with angle of attack was 
primarily due to increase in base pressure force. 



NACA RM E51H23 

WIRE CLOTH AS POROUS MATERIAL FOR 
TRANSPIRATION-COOLED WALLS. E. R. G. 
Eckert, Martin R. Kinsler, and Reeves P. Cochran. 
November 1951. 38p. diagrs., photos., tab. (NACA 
RM E51H23) (Declassified from Confidential, 
9/15/55) 

The permeability characteristics and tensile strength 
of a porous material developed from stainless-steel 
corduroy wire cloth for use in transpiration-cooled 
vvalls where the primary stresses are in one direc- 
tion were investigated. The results of this investi- 
gation are presented and compared with similar re- 
sults obtained with porous sintered metal compacts. 
A much wider range of permeabilities is obtainable 
with the wire cloth than with the porous metal com- 
pacts considered and the ultimate tensile strength in 
the direction of the primary stresses for porous ma- 
terials produced from three mesh sizes of wire cloth 
is from two to three times the ultimate tensile 
strengths of the porous metal compacts. 



NACA RM E52I12 

INTERSTAGE SURVEYS AND ANALYSIS OF VIS- 
COUS ACTION IN LATTER STAGES OF A MULTI- 
STAGE AXIAL-FLOW COMPRESSOR. William B. 
Briggs and Charles C. Giamati. March 1953. 51p. 
diagrs., photo., tab. (NACA RM E52I12) 
(Declassified from Confidential, 9/15/55) 

The overall performance of an eight-stage axial- 
flow compressor having a design stage pressure 
ratio of 1.23 was determined and radial interstage 
surveys were made for four weight flows at both 50 
percent and 100 percent design speed. The survey 
data, from behind the guide vanes and each row of 



NACA 

RESEARCH ABSTRACTS NO. 91 



the fifth to eighth stages, were presented as axial 
and tangential velocity components from which 
boundary-layer parameters of displacement and 
momentum thickness were calculated. An analysis 
related current qualitative ideas of flows which are 
imposed upon the through flow and the quantitative 
status of theory to the observed variation of these 
parameters. Midchannel velocity peaks in excess 
of design value were not found. 



NACA RM E51I25 

FLOW SEPARATION AHEAD OF A BLUNT AXDALLY 
SYMMETRIC BODY AT MACH NUMBERS 1.76 TO 
2.10. W. E. Moeckel. December 1951. 12p. 
diagrs., photos. (NACA RM E51I25) (Declassified 
from Confidential, 9/15/55) 

The pressure distribution and drag were determined 
for a spherical-nosed axially symmetric body with 
thin projecting rods at Mach numbers of 1.76, 1.93, 
and 2.10. The upstream projection distance of the 
rods was varied over a wide range to study changes 
in the character of the flow separation and to deter- 
mine the variation of drag and pressure distribution 
with tip projection. Drag coefficients between 0.18 
and 0.30 were obtained for most tip projections at 
each Mach number. 



NACA RM L8I08 

LONGITUDINAL-STABILITY INVESTIGATION OF 
HIGH-LIFT AND STALL-CONTROL DEVICES ON A 
52° SWEPTBACK WING WITH AND WITHOUT FUSE- 
LAGE AND HORIZONTAL TAIL AT A REYNOLDS 
NUMBER OF 6.8 x 10 6 . Gerald V. Foster and 
James E. Fitzpatrick. December 20, 1948. 41p. 
diagrs., photos., tabs. (NACA RM L8I08) 
(Declassified from Confidential, 9/15/55) 

Contains low-speed longitudinal stability character- 
istics of a 52° sweptback wing of aspect ratio 2.88, 
taper ratio 0.625, and NACA 64j-112 airfoil sections 
normal to the 0.282-chord line, in combination with 
split flaps, leading-edge flaps, and upper-surface 
fences. Low-wing and midwing-fuselage aerody- 
namic characteristics are presented with and without 
a horizontal tail at various vertical locations. Tests 
were conducted at a Reynolds number of 6.8 x 10". 



NACA RM L8I30a 

EFFECTS OF A SWEPTBACK HYDROFOIL ON THE 
FORCE AND LONGITUDINAL STABILITY CHARAC- 
TERISTICS OF A TYPICAL HIGH-SPEED AIR- 
PLANE. Raymond B. Wood. December 2, 1948. 
19p. diagrs., photo., tabs. (NACA RM L8I30a) 
(Declassified from Confidential, 9/15/55) 

An investigation was conducted in the Langley 8-foot 
high-speed tunnel to determine the effects of a swept- 
back hydrofoil on the force and longitudinal stability 



NACA 

RESEARCH ABSTRACTS NO. 9) 



characteristics of a typical high-speed airplane. 
The Mach number range for this investigation was 
from 0.60 to 0.95 and at M = 1.20. The effects of 
the hydrofoil on the lift, drag, and pitching-moment 
characteristics are presented. 



NACA RM L9I30 

FLIGHT INVESTIGATIONS AT HIGH-SUBSONIC, 
TRANSONIC, AND SUPERSONIC SPEEDS TO DE- 
TERMINE ZERO-LIFT DRAG OF FIN-STABILIZED 
BODIES OF REVOLUTION HAVING FINENESS 
RATIOS OF 12.5, 8.91, AND 6.04 AND VARYING 
POSITIONS OF MAXIMUM DIAMETER. Roger G. 
Hart and Ellis R. Katz. November 30, 1949. 36p. 
diagrs., photos. (NACA RM L9I30) 
(Declassified from Confidential, 9/15/55) 

Rocket-powered models were flown at transonic and 
supersonic speeds to determine the zero-lift drag of 
fin-stabilized bodies of revolution differing in fine- 
ness ratio and in position of maximum diameter. 
The bodies were of fineness ratio 12.5, 8.91, and 
6.04 and all had cut-off sterns with equal base and 
frontal areas. 



NACA RM L50I27 

THE LONGITUDINAL STABILITY, CONTROL EF- 
FECTIVENESS, AND CONTROL HINGE-MOMENT 
CHARACTERISTICS OBTAINED FROM A FLIGHT 
INVESTIGATION OF A CANARD MISSILE CONFIG- 
URATION AT TRANSONIC AND SUPERSONIC 
Sl'EEDS. Roy J. Niewald and Martin T. Moul. 
November 24, 1950. 43p. diagrs., photos. 
(NACA RM L50I27) 
(Declassified from Confidential, 9/15/55) 

A 60° delta-wing canard missile configuration was 
flight-tested at the Langley Pilotless Aircraft 
Research Station at Wallops Island, Va. Longitudi- 
nal stability derivatives, control hinge -moment, and 
drag characteristics were obtained at transonic and 
supersonic velocities by utilizing a continuous step 
control disturbance of ±5°. 



NACA RM L51G31 

SYSTEMATIC TWO-DIMENSIONAL CASCADE TESTS 
OF NACA 65-SERIES COMPRESSOR BLADES AT 
LOW SPEEDS. L. Joseph Herrig, James C. Emery, 
and John R. Erwin. September 1951. 223p. 
diagrs., photo., tabs. (NACA RM L51G31) 
(Declassified from Confidential, 9/15/55) 

A two-dimensional low-speed porous-wall cascade 
tunnel investigation has been conducted to establish 
the performance of the NACA 65-series compressor 
blade sections over the useful range of inlet angle, 
solidity, and section camber. Design points for opti- 
mum high-speed operation are presented. The load- 
ing limitation is determined for some conditions. 
Trends of section operating range with increasing 
section camber are determined for the four inlet 
angles tested. 



NACA RM L51H03 

ROLLING EFFECTIVENESS OF ALL-MOVABLE 
WINGS AT SMALL ANGLES OF INCIDENCE AT 
MACH NUMBERS FROM 0.6 TO 1.6. H. Kurt 
Strass and Edward T. Marley. October 1951. 16p. 
diagrs., photo., tab. (NACA RM L51H03) 
(Declassified from Confidential, 9/15/55) 

An investigation of the rolling effectiveness of sev- 
eral all-movable wing configurations has been con- 
ducted throughout a Mach number range from 0.6 to 
1.6 in order to check a simplified wing incidence cor- 
rection theory which states that for all-movable wing 

pb 2i w (\ + 2x\ , .... 

5^ = * - h o ( i oy )• In addition, a comparison was 

made with two other more complex methods of esti- 
mation. The results showed that the simplified 
theory gave accurate agreement with experiment at 
Mach numbers from 0.6 to 1.6 and suggest the use of 
this simplified equation as a means of predicting the 
rolling effectiveness of all-movable wings throughout 
this speed range. 



NACA RM L52I16 

SOME EXPERIMENTAL STUDIES OF PANEL 
FLUTTER AT MACH NUMBER 1.3. Maurice A. 
Sylvester and John E. Baker. December 1952. 
25p. diagrs., photos., tab. (NACA RM L52I16) 
(Declassified from Confidential, 9/15/55) 

Experimental studies of panel flutter using thin metal 
plates .vere conducted at a Mach number of 1.3 to 
verify its existence and to study the effects of some 
structural parameters on the flutter characteristics. 
The effects of tensile forces and buckling were 
studied on panels clamped front and rear, in addition 
to initially buckled panels clamped on all four edges. 
Panel flutter was obtained under these laboratory 
conditions and it was found that tensile forces, 
shortening the panels, and increasing the bending 
stiffness were effective means for eliminating flutter. 
Buckled panels were more susceptible to flutter than 
unbuckled panels. No apparent systematic trends in 
the flutter modes or frequencies could be observed. 



NACA RM L53F17 

METHOD OF ESTIMATING THE INCOMPRESSIBLE- 
FLOW PRESSURE DISTRIBUTION OF COMPRESSOR 
BLADE SECTIONS AT DESIGN ANGLE OF ATTACK. 
John R. Erwin and Laura A. Yacobi. December 
1953. 41p. diagrs., tab. (NACA RM L53F17) 
(Declassified from Confidential, 9/15/55) 

A method was devised for estimating the incom- 
pressible flow pressure distribution over com- 
pressor blade sections at design angle of attack. 
The theoretical incremental velocities due to camber 
and thickness of the section as an isolated airfoil 
are assumed proportional to the average passage 
velocity and are modified by empirically determined 
interference factors. Comparisons were made be- 
tween estimated and test pressure distributions of 
NACA 65-series sections for typical conditions. 
Good agreement was obtained. 



10 



NACA RM L53I11 



UNIVERSITY OF FLORIDA 



3 1262 08153 283 9 



NACA 
RESEARCH 



ABSTRACTS NO. 91 



FREE-FLIGHT -TUNNEL INVESTIGATION OF THE 
LOW-SPEED STABILITY AND CONTROL CHARAC- 
TERISTICS OF A CANARD AIRPLANE MODEL. 
Joseph L. Johnson, Jr., and John W. Paulson. 
October 1953. 37p. diagrs., photo., 2 tabs. (NACA 
RM L53I11) (Declassified from Confidential, 
9/15/55) 

Results are presented of an experimental investiga- 
tion in the Langley free-flight tunnel to determine 
the dynamic lateral stability and control character- 
istics of a model of a canard-type airplane. Tests 
were made with several vertical-tail configurations 
for the model with a triangular horizontal tail and 
with a sweptback horizontal tail having a leading- 
edge flap. 



NACA RM L53I25a 

FLIGHT DETERMINATION OF DRAG OF NORMAL- 
SHOCK NOSE INLETS WITH VARIOUS COWLING 
PROFILES AT MACH NUMBERS FROM 0.9 TO 1.5. 
R. I. Sears, C. F. Merlet, and L. W. Putland. 
October 1953. 36p. diagrs., photos., tabs. (NACA 
RM L53I25a) (Declassified from Confidential, 
9/15/55) 

External-drag data are presented for normal-shock 
nose inlets with 1 -series, parabolic, and conic 
cowling profiles. The tests were made at an angle 
of attack of 0° by using rocket -propelled models in 
free flight at Mach numbers from 0.9 to 1.5. The 
Reynolds number based on body maximum diameter 
varied from 2.5 x 10° to 5.5 x 10°\ At maximum 
flow rate, the inlet models had about the same ex- 
ternal drag at a Mach number of approximately 1.1, 
but at higher Mach numbers the sharp-lip conic cowl 
had the least drag. Blunting or beveling the lip of 
the conic cowl while keeping the fineness ratio 
constant resulted in drag coefficients slightly higher 
than for the sharp-lip conic cowl at maximum flow 
rate. At a mass-flow ratio of about 0.8, the conic 
cowls with sharp, blunt, or beveled lips and the 
parabolic cowl all gave about the same drag. 



NACA RM L53I29b 



INVESTIGATION AT SUPERSONIC SPEEDS OF THE 
VARIATION WITH REYNOLDS NUMBER AND MACH 
NUMBER OF THE TOTAL, BASE, AND SKIN- 
FRICTION DRAG OF SEVEN BOATTAIL BODIES OF 
REVOLUTION DESIGNED FOR MINIMUM WAVE 
DRAG. August F. Bromm, Jr.. and Julia M. 
Goodwin. December 1953. 20p. diagrs., photo. 
(NACA RM L53I29b) (Declassified from 
Confidential, 9/15/55) 

Results are presented from an investigation of the 
variation with Reynolds number and Mach number of 
the total, base, and skin-friction drag of seven boat- 
tail bodies of revolution designed for minimum wave 
drag according to the theory of NACA TN 2550. The 
tests covered a Reynolds number range from approxi- 
mately 1.0 x 10 6 to 10.0 x 10 6 at Mach numbers of 
1.62, 1.93, and 2.41, respectively. 



NACA RM L53I30b 

TWO-DIMENSIONAL LOW -SPEED CASCADE INVES- 
TIGATION OF NACA COMPRESSOR BLADE SEC- 
TIONS HAVING A SYSTEMATIC VARIATION IN 
MEAN-LINE LOADING. John R. Erwin, Melvyn 
Savage, and James C. Emery. November 1953. 
129p. diagrs., tabs. (NACA RM L53I30b) 
(Declassified from Confidential, 9/15/55) 

The low-speed cascade performance of the high- 
speed NACA 65- (C-i A2l8b) 1( ^ compressor blade 

sections has been systematically investigated. When 
used in conjunction with published cascade data, the 
results will provide design information for all inlet 
angle and solidity conditions within the usual range of 
application. Summary curves have been prepared to 
facilitate the selection of blade sections and settings 
to fulfill the conditions dictated by compressor design 
velocity diagrams. Comparative tests of blade 
sections having widely different loading distributions 
indicated that these data, in conjunction with previ- 
ously published cascade data, permit a fairly 
accurate prediction of design performance for most 
compressor blade sections since the mean lines 
tested probably encompass the practical range of 
compressor-blade mean-line loading distributions. 
A comparative evaluation of the high-speed per- 
formance capabilities of the blade sections investi- 
gated was made. 



NACA - Langley Field, Va.